This invention relates generally to gas turbine engine airfoils, and more particularly to apparatus and methods for cooling hollow turbine airfoils.
A typical gas turbine engine includes a turbomachinery core having a high pressure compressor, a combustor, and a high pressure turbine in serial flow relationship. The core is operable in a known manner to generate a primary gas flow. The high pressure turbine (or “HPT”) includes one or more stages which extract energy from the primary gas flow. Each stage comprises row of stationary vanes or nozzles that direct gas flow into a downstream row of blades or buckets carried by a rotating disk. These components operate in an extremely high temperature environment. To ensure adequate service life, the vanes and blades are hollow and are provided with a flow of coolant, such as air extracted (bled) from the compressor. This coolant flow is circulated through the hollow airfoil's internal coolant path and is then exhausted through a plurality of cooling holes.
One type of cooling hole that has been found effective is a shaped or diffuser hole that includes a circular metering portion and a flared portion that acts as a diffuser. The shaped diffuser holes can be oriented axially or parallel to the gas stream (indicated by the arrow “G” in FIG. 1), or they can be oriented vertically at various angles relative to a radial line drawn to engine centerline. Recent experience with HPT airfoils has shown that reduced airfoil casting wall thickness because of manufacturing process variation can reduce diffuser hole effectiveness. This can be countered by increasing wall thickness for the entire airfoil, but this results in undesirable weight increase.
Accordingly, there is a need for a turbine airfoil with diffuser holes that perform effectively without excessive weight increase.